1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled blade in a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine with multiple rows or stages of rotor blades that react with a high temperature gas flow to drive the engine or, in the case of an industrial gas turbine (IGT), drive an electric generator and produce electric power. It is well known that the efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage vanes and blades and the amount of cooling that can be achieved for these airfoils.
In latter stages of the turbine, the gas flow temperature is lower and thus the airfoils do not require as much cooling flow. In future engines, especially IGT engines, the turbine inlet temperature will increase and result in the latter stage airfoils to be exposed to higher temperatures. To improve efficiency of the engine, low cooling flow airfoils are being studied that will use less cooling air while maintaining the metal temperature of the airfoils within acceptable limits. Also, as the TBC (thermal barrier coating) gets thicker, less cooling air is required to provide the same metal temperature as would be for a thicker TBC.
FIG. 1 shows a prior art turbine rotor blade with a 1+3 serpentine flow cooling circuit for the blade mid-chord serpentine cooling. The airfoil leading edge is cooled with a backside impingement cooling in conjunction with leading edge showerhead film cooling holes 11 and pressure side 12 and suction side 13 gill holes. Cooling air for the leading edge region is supplied through a separate radial supply channel 14 through a row of metering and impingement holes 15. FIG. 2 shows a flow diagram of the blade cooling circuit of FIG. 1. The airfoil main body is cooled with a triple pass (also referred to as a 3-pass) forward flowing serpentine circuit with a cooling air supply channel being the first leg 21, a second leg 22 and a third leg 23 in conjunction with pressure side 25 and suction 26 side film cooling holes and trailing edge discharge cooling holes 27. Blade tip cooling holes 28 are also used in both the leading edge cooling supply channel 14 and the 3-pass serpentine flow circuit to discharge some of the cooling air through the blade tip.
In the prior art, blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine coolant passages from both of the pressure and suction side surfaces near to the blade tip edge and the top surface of the squealer cavity or pocket. In general, film cooling holes are formed along the airfoil pressure side and suction side tip sections, from the leading edge to the trailing edge in order to provide edge cooling for the blade squealer tip. Also, convective cooling holes are also formed along the tip rail on the inner surface of the squealer pocket to provide additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow field of hot gas flow, a large quantity of film cooling holes and cooling flow is required for cooling of the blade tip periphery. FIG. 3 shows a prior art blade with the tip edge film cooling holes 31 on the pressure side wall of the blade and FIG. 4 shows a detailed view of the film cooling hole 31 with its breakout shape. FIG. 5 shows the prior art blade with film cooling holes 32 on the suction side wall adjacent to the tip edge and FIG. 6 shows the breakout shape for the film hole 32. The hot gas vortex flow 35 is shown in FIG. 5 forming along the suction side wall at the trailing edge region of the blade.
For the prior art FIGS. 1 and 2 design, the last leg 23 of the 3-pass serpentine flow circuit is determined by the ceramic core manufacturing requirements. As a result of this cooling flow design requirement, when the cooling air is bled off from the cavity for cooling of both the pressure side and suction side walls and along the blade tip section, the spanwise internal Mach number of the cooling air flow becomes very low. A high Mach number for the cooling air flow will produce high heat transfers from the hot metal surface to the cooling air. This results in a lower flow through velocity and a cooling side internal heat transfer coefficient. The same flow phenomena can also be applied to the airfoil leading edge cooling supply channel.